A gas turbine engine includes in serial flow communication, one or more compressors followed in turn by a combustor and high and low pressure turbines, disposed symmetrically about a longitudinal axis centerline within an annular outer casing.
Each of the turbines includes one or more stages of rotor blades extending radially outwardly from respective rotor disks, with the blade tips being disposed closely adjacent to a turbine shroud assembly supported within the casing. It is desirable to maintain the gap between the blade tips and the shroud assembly as small as possible throughout the engine operation range because the combustion gas flowing therethrough bypasses the turbine blades and therefore provides no useful contribution. However, because the material of the stator components and the turbine rotor are different, and because inertia has an influence on the expansion of the rotor, the stator components, i.e. the engine case, outer air seal, and support mechanism, expand at a different rate than the expansion of the rotor. Therefore, the gap must be sized larger than would otherwise be desirable.
Conventionally, small gas turbine engines typically use a passive tip clearance control system when attempts are made to optimize the thermal response characteristics of the rotor and the casing. Full pressure compressor air is used both as the cooling medium and as the air seals around the blade tips, and is then exhausted into the turbine combustion gas path. When operating the engine during a transitional period, the thermal response rates of the casing and the rotor blades are difficult to match, thereby resulting in a pinch-point. This pinch-point causes a system limitation as to the minimum achievable tip clearance without rubbing.
Larger engines usually use active tip clearance control where inter-stage compressor bleed air is used to externally cool the turbine casing, typically in an impingement manner. This inter-stage compressor bleed air can be turned off by a valve during initial operation so as to avoid the pinch-point. When the engine has thermally stabilized, the valve is opened and the turbine casing effectively contracts to minimize tip clearance. Typically, this inter-stage compressor bleed air is dumped into the nacelle and lost to the cycle after having cooled the turbine casing.
Various efforts have been made to improve turbine tip clearance control in gas turbine engines. Examples of those efforts are illustrated in U.S. Pat. No. 4,513,5697 to Deveau et al. on Apr. 30, 1985; U.S. Pat. Nos. 5,593,277, 5,562,408 and 5,553,999 issued to Proctor et al. on Jan. 14, 1997, Oct. 8, 1996 and Sep. 10, 1996 respectively; U.S. Pat. No. 5,048,288, issued to Bessette et al. on Sep. 17, 1991; U.S. Pat. No. 4,358,926, issued to Smith on Nov. 16, 1982; and U.S. Pat. No. 6,487,491, issued to Karpman et al. on Nov. 26, 2002. These prior art patents disclose method, systems and apparatuses for improving turbine tip clearance control in one or more aspects of this matter. Nevertheless, continuous efforts to develop the technology in this field are still needed in order to achieve better performance of gas turbine engines, particularly for use with aircraft. The prior art offers complex solutions and solutions which do not maximize the efficiency of cooling air systems in the engine. Improvements are therefore desired.